Spacecraft generally use solar cells to collect solar radiation and convert it into the electrical power necessary to operate the spacecraft. The solar cells are normally disposed on a solar array. A solar array typically comprises one or more solar panels electrically attached to each other and to the spacecraft. Each solar panel in an array typically comprises numerous individual solar cells, which are usually laid out in rows and connected together electrically at their adjacent edges. These photocells form a two-dimensional array and are frequently mounted on a solar panel comprising lightweight graphite face sheets covering a honeycomb core. When multiple solar panels are connected together to power the spacecraft, these panels must fold up, typically accordion style, prior to launch and unfold once in orbit.
The number of solar cells that must be employed is a function of the anticipated spacecraft power demand and the efficiency of the cells. High-efficiency solar cells are typically employed to reduce the area of photovoltaics required by a specific spacecraft. This reduces panel area and thus overall mass from the required supporting structure and minimizes the volume of the stowed power system. But such cell devices are quite expensive. Because system cost and mass both increase directly with the number of solar cells employed, there is considerable incentive to reduce the quantity of solar cells that a spacecraft must carry on an array.
The efficiency of solar panels for spacecraft is generally evaluated on the basis of numerous criteria, including watts per kilogram, watts per cubic meter (stowed), and watts per dollar. Solar cells are the most expensive component of a solar array. To reduce solar array cost, solar concentrators may be used to reduce the number of cells. Lightweight reflective surfaces have been used in various combinations with known solar panels to improve their efficiency. Examples of such combinations are disclosed in U.S. Pat. No. 6,017,002 to Burke et al. for Thin-Film Solar Reflectors Deployable From An Edge-Stowed Configuration, U.S. Pat. No. 5,520,747 to Marks for a Foldable Low Concentration Solar Array, and U.S. Pat. No. 5,885,367 to Brown et al. for a Retractable Thin Film Solar Concentrator For Spacecraft, which all use large deployable sheets to reflect onto the solar panel as depicted in FIG. 8 herein. These patents are all incorporated herein by reference as if set forth in their entirety. As can be seen, use of these so called “full panel reflector systems” greatly increases the collection area of the solar panel. One considerable drawback of such full panel reflector systems, however, is that because the collection area of the solar panel is large in comparison to the panel's radiative area, the solar cells operate at high temperature, which reduces their efficiency. Another potential drawback of full panel reflector systems is that because the reflectors are large in comparison to the area of an individual solar cell, relatively minor distortions in the reflectors can cause significant differentials in the amount of solar energy reflected onto each cell, which in turn can impair the efficiency of the panel. Similarly, because the reflector in full panel reflector systems is far away from at least some of the solar cells on the panel, small distortions in the reflector can cause significant differentials in the amount of light reflected onto the individual solar cells.
As noted above, the harvesting of a large cross-sectional area of solar energy from a smaller area of solar cells by concentration is a well-recognized art for spacecraft and has been used to improve the efficiency and other performance parameters of known solar panels. Numerous other techniques are employed terrestrially, including using lenses such as Fresnel lenses to refract the energy onto the cells, and large mirror arrangements to reflect the energy onto the cells. However effective these known devices may be in directing energy from a larger area onto a smaller area, they bring with them many problems of practical concern when used for spacecraft.
Designing land-based apparatus to capture solar energy involves fewer constraints. The apparatus is built in place and stays there. It need not automatically deploy into configuration. Weight is no concern, and neither is perfect reliability, because within reason, the apparatus is accessible and readily repaired. Structural efficiency is really not an issue; a ground-based device may simply be made as heavy and strong as desired, with a generous allowance for safety. Thus, the weight, reliability, and rigidity of the apparatus do not impose any special concerns for the design of land-based solar energy systems. Nor does the variability of environmental conditions such as temperature create any special design concerns. In land-based solar energy systems, nearly all design concerns can be minimized or corrected by over-design of the apparatus.
Such is not the circumstance for spacecraft. Weight is a primary consideration, not only because of the cost per kilogram to launch the apparatus, but because weight of one part of a spacecraft will necessarily require a reduction of weight elsewhere due to the ultimate limitation on the total launch weight of the entire craft.
Reliability is also a prime concern. Spacecraft are one-way vehicles. Once in space, they remain there during their useful life, and except in a few extraordinary situations such as the Hubble Telescope, they will never be approached after launch. The failure of apparatus such as a solar array dooms all or a large part of the intended life and function of the entire craft.
Rigidity in the sense of maintenance of shape under varying conditions is made complicated by the extreme variations in temperature as the apparatus enters and leaves the shadow of the earth. While in the shadow, temperatures as low as −180° C. are endured. While out of the shadow and exposed directly to the sun, temperatures as high as 110° C. are endured. When the solar panels transition between light and shadow, the change in temperature of the apparatus occurs in only a few minutes, and does not occur uniformly throughout. This results in a reaction known as “thermal snap” in which the distortions that result from rapid temperature change cause a quick bending distortion that shudders the spacecraft and can damage the array. As the wing temperatures change over the sunlit portion of the orbit, distortions of the structure can cause the concentrator optics to malfunction.
High temperatures are also the enemy of solar cells. The efficiency of solar cells decreases as their temperature increases. It is, therefore, important to mount the cells in an arrangement such that the energy received by them does not heat the cells to an unacceptable temperature.
This is further complicated if large reflecting areas are involved where there may be localized higher temperatures due to distortions of the reflector. The overheated cells will function less efficiently. Even isolated instances of under-performing solar cells can impair the efficiency of the solar panel. The problems created by under-performing solar cells are exacerbated by the fact that the solar cells within individual rows in known solar arrays are electrically connected in series, which means that the electrical output of an entire row of cells will be compromised by even one cell's under-performance.
Thus, it can be seen that there is a daunting array of considerations in the design of solar energy systems for spacecraft. Various arrangements have been proposed in the past for improving the efficiency and resistance to environmental and hostile threats of solar energy systems. Over the decades, there has been a long succession of solar arrays produced and launched. Many have been successful, but a disheartening proportion of them have failed partially or totally, causing the loss of very costly space vehicles, or a major reduction in their useful life.
A need, therefore, exists for a solar cell array configuration that optimizes solar collection without risking non-uniform cell illumination or unacceptably increasing the temperature of the solar cells and thereby impairing their efficiency. A need further exists for such a system that is simple, lightweight, and reliable in deployment and operation.